1. Field of the Invention
The present invention relates to a cooled shroud having a cooling structure in a gas turbine stationary blade.
2. Description of the Prior Art
FIG. 5 is a perspective view showing one typical example of a cooling system in a prior art gas turbine stationary blade. In the figure, numeral 30 designates a stationary blade, numeral 31 designates an outer shroud thereof and numeral 32 designates an inner shroud. Numeral 33, 34 and 35, respectively, designates an insert, arranged in the order of place from a leading edge side toward a trailing edge side, inserted in the direction from the outer shroud 31 toward the inner shroud 32, numeral 33a, 34a and 35a, respectively, designates an air injection hole provided to the respective insert and numeral 36 designates a trailing edge fin. Numeral 37 generally designates an air injection hole provided in a blade surface for blowing air, wherein numerals 37a and 37c are shown in the figure and numeral 37b is not shown.
In the stationary blade 30 of the above structure, cooling air is introduced through the outer shroud 31 and the inner shroud 32, respectively, into the inserts 33, 34 and 35 and is blown from the air injection holes 33a, 34a and 35a toward a blade inner surface to perform an impingement cooling of the blade inner surface and is then blown from the air injection holes 37a, 37b and 37c, provided in the blade surface, to perform a shower head cooling, a film cooling and a pin fin cooling of the blade.
Also, as for the cooling of the shroud, it is done, as shown in the figure as one example, such that the cooling air which flows in is injected to impinge rectangularly on an impingement plate 39 which is provided in the inner shroud 32 in parallel thereto so that the air passes through a multiplicity of holes to be diffused to cool an entire surface of the inner shroud 32 and is then flown out of a rear end of the shroud.
FIG. 6 shows another example of a stationary blade cooling system. In the figure, numeral 40 designates a stationary blade, numeral 41 designates an outer shroud thereof and numeral 42 designates an inner shroud. Numeral 43A, 43B, 43C, 43D and 43E, respectively, designates an air passage, numeral 45 designates an air injection hole of a trailing edge and numeral 46 designates a turbulator provided to an inner wall of the air passage 43A to 43D, respectively, for making the air flow there turbulent for enhancement of a heat transfer.
In the present stationary blade cooling system, a cooling air 47 flows in from the outer shroud 41 into the air passage 43A to flow to a base portion, to enter therefrom the next air passage 43B, to flow to a tip portion, to enter therefrom the next air passage 43C and then likewise to flow in the air passages 43D and 43E, sequentially, to cool the blade, and is blown out of the air passage 43E through the air injection hole 44 of the trailing edge and a remaining air flows down out of the inner shroud 42.
In the cooling system shown in FIG. 6, while there is constructed a serpentine cooling passage by the air passages 43A to 43E so that air flows there to cool the blade, there is considered nothing of a cooling of the shroud.
In the prior art gas turbine cooling system as described above, the cooling of the shroud is done, as the example shown in FIG. 5, such that the cooling air is flown to impinge on the impingement plate provided in the shroud and is flown in the shroud through the multiplicity of holes to cool the shroud and then is discharged through an air passage in the rear end of the shroud, however, there is applied no cooling of the shroud in the stationary blade of a turbine rear stage side, as shown in FIG. 6.
The cooling of the shroud by use of such an impingement plate as mentioned above is not necessarily sufficient and, moreover, there are many cases where such cooling is done on the shroud on a turbine front stage side but no cooling is done on the turbine rear stage side, hence, a means for further enhancing the cooling effect as a whole has been desired.